Method and apparatus for handling pre-diffuser airflow for cooling high pressure turbine components

ABSTRACT

A gas turbine engine is provided comprising a compressor section, a combustor section, a diffuser case module, and a manifold. The diffuser case module includes a multiple of struts within an annular flow path from said compressor section to said combustor section, wherein at least one of said multiple of struts defines a mid-span pre-diffuser inlet in communication with said annular flow path. The manifold is in communication with said mid-span pre-diffuser inlet and said compressor section.

This application claims priority to PCT Patent Application No.PCT/US14/19585 filed Feb. 28, 2014, which claims priority to U.S. PatentAppln. No. 61/770,853 filed Feb. 28, 2013.

BACKGROUND

The present disclosure relates to a gas turbine engine and, moreparticularly, to a cooling architecture therefor.

Gas turbine engines, such as those that power modern commercial andmilitary aircraft, generally include a compressor section to pressurizean airflow, a combustor section to burn a hydrocarbon fuel in thepresence of the pressurized air, and a turbine section to extract energyfrom the resultant combustion gases.

Thermal loads within the gas turbine engine vary. Such variance mayaffect performance even within the bounds of material specifications.

SUMMARY

A diffuser for a gas turbine engine according to one disclosednon-limiting embodiment of the present disclosure includes an outershroud, an inner shroud, a multiple of struts between said outer shroudand said inner shroud to define an annular flow path, and a mid-spanpre-diffuser inlet in communication with said annular flow path.

In a further embodiment of the foregoing embodiment, the mid-spanpre-diffuser inlet includes an inlet directed generally parallel to acore airflow thru said annular flow path.

In a further embodiment of any of the foregoing embodiments, themid-span pre-diffuser inlet includes an inlet on both sides of an outerairfoil wall surface of at least one of said multiple of struts.

In a further embodiment of any of the foregoing embodiments, themid-span pre-diffuser inlet includes an inlet on one side of an outerairfoil wall surface of at least one of said multiple of struts.

In a further embodiment of any of the foregoing embodiments, themid-span pre-diffuser inlet is a NACA inlet. In the alternative oradditionally thereto, in the foregoing embodiment the NACA inletprovides a capacity of approximately 0%-2.5% of a core airflow.

In a further embodiment of any of the foregoing embodiments, themid-span pre-diffuser inlet is a ram inlet. In the alternative oradditionally thereto, in the foregoing embodiment the ram inlet providesa capacity of approximately 2.5%-5% of a core airflow.

In a further embodiment of any of the foregoing embodiments, themid-span pre-diffuser inlet is an annular inlet. In the alternative oradditionally thereto, in the foregoing embodiment the annular inletprovides a capacity of approximately 10%-20% of a core airflow. In thealternative or additionally thereto, in the foregoing embodiment theannular inlet is located circumferentially between each of said multipleof struts.

In a further embodiment of any of the foregoing embodiments, thediffuser includes a manifold in communication with said mid-spanpre-diffuser inlet. In the alternative or additionally thereto, in theforegoing embodiment the manifold communicates a temperature tailoredairflow.

A gas turbine engine according to another disclosed non-limitingembodiment of the present disclosure includes a compressor section, acombustor section, a diffuser case module with a multiple of strutswithin an annular flow path from said compressor section to saidcombustor section, at least one of said multiple of struts defines amid-span pre-diffuser inlet in communication with said annular flowpath, and a manifold in communication with said mid-span pre-diffuserinlet.

In a further embodiment of the foregoing embodiment, the manifold is incommunication with a region of the gas turbine engine to supply atemperature tailored airflow thereto.

A method of communicating an airflow within a gas turbine engineaccording to another disclosed non-limiting embodiment of the presentdisclosure includes tapping a pre-diffuser airflow.

In a further embodiment of the foregoing embodiment, the method includescommunicating the pre-diffuser airflow thru at least one of a multipleof struts.

In a further embodiment of any of the foregoing embodiments, the methodincludes communicating the pre-diffuser airflow thru at least one of amultiple of struts at a radial mid-strut location.

In a further embodiment of any of the foregoing embodiments, the methodincludes communicating the pre-diffuser airflow thru an annular inletwhich extends between a multiple of struts.

In a further embodiment of any of the foregoing embodiments, the methodincludes tapping between 0%-20% of a compressor section airflow.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled in the art fromthe following detailed description of the disclosed non-limitingembodiment. The drawings that accompany the detailed description can bebriefly described as follows:

FIG. 1 is a schematic cross-section of an example gas turbine enginearchitecture;

FIG. 2 is a partial expanded cross-section view of a hot section of agas turbine engine shown in FIG. 1;

FIG. 3 is an expanded front view of a diffuser case of the combustorsection;

FIG. 4 is an expanded sectional view of a pre-diffuser strut accordingto one disclosed non-limiting embodiment;

FIG. 5 is a cross-sectional view of an inner diffuser case with amanifold in communication with a pre-diffuser strut according to onedisclosed non-limiting embodiment;

FIG. 6 is a radial cross-section of the inner diffuser illustrating aninlet according to one disclosed non-limiting embodiment;

FIG. 7 is a cross-section thru the inlet strut of FIG. 6 along line 7-7;

FIG. 8 is a radial cross-section of the inner diffuser illustrating aninlet according to another disclosed non-limiting embodiment;

FIG. 9 is a cross-section thru the inlet strut of FIG. 8 taken alongline 9-9;

FIG. 10 is a radial cross-section of the inner diffuser of FIG. 11;

FIG. 11 is a front view of the inner diffuser of FIG. 10

FIG. 12 is a schematic view of a gas turbine engine hot sectionillustrating an airflow communication scheme for fuel-nozzlepre-swirlers according to one disclosed non-limiting embodiment;

FIG. 13 is a schematic view of a gas turbine engine hot sectionillustrating an airflow communication scheme for a high pressurecompressor (HPC) blade attachment hardware according to anotherdisclosed non-limiting embodiment;

FIG. 14 is a schematic view of a gas turbine engine hot sectionillustrating an airflow communication scheme for HPC blade attachmenthardware with a heat exchanger according to another disclosednon-limiting embodiment;

FIG. 15 is a schematic view of a gas turbine engine hot sectionillustrating an airflow communication scheme for HPC blade attachmenthardware with a heat exchanger and buffer system according to anotherdisclosed non-limiting embodiment;

FIG. 16 is a radial cross-section of a strut with a buffer airpassageway;

FIG. 17 is a schematic view of a gas turbine engine hot sectionillustrating an airflow communication scheme for HPC blade attachmenthardware with a heat exchanger and buffer system according to anotherdisclosed non-limiting embodiment;

FIG. 18 is a schematic view of a gas turbine engine hot sectionillustrating an airflow communication scheme for a combustor with a heatexchanger and buffer system according to another disclosed non-limitingembodiment;

FIG. 19 is a schematic view of a gas turbine engine hot sectionillustrating an airflow communication scheme for HPT first vanes with aheat exchanger and buffer system according to another disclosednon-limiting embodiment;

FIG. 20 is a schematic view of a gas turbine engine hot sectionillustrating an airflow communication scheme for a BOAS system with aheat exchanger and buffer system according to another disclosednon-limiting embodiment;

FIG. 21 is a schematic view of a gas turbine engine hot sectionillustrating an airflow communication scheme for HPT blade attachmenthardware with a heat exchanger and buffer system according to anotherdisclosed non-limiting embodiment;

FIG. 22 is a schematic view of a gas turbine engine hot sectionillustrating an airflow communication scheme for HPT second stage vaneswith a heat exchanger and buffer system according to another disclosednon-limiting embodiment; and

FIG. 23 is a schematic view of a gas turbine engine hot sectionillustrating an airflow communication scheme for HPT inter-stage sealhardware with a heat exchanger and buffer system according to anotherdisclosed non-limiting embodiment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative enginearchitectures might include an augmentor section and exhaust ductsection (not shown) among other systems or features. The fan section 22drives air along a bypass flowpath while the compressor section 24drives air along a core flowpath for compression and communication intothe combustor section 26 then expansion thru the turbine section 28.Although depicted as a turbofan in the disclosed non-limitingembodiment, it should be understood that the concepts described hereinare not limited to use with turbofans as the teachings may be applied toother types of turbine engines such as a low bypass augmented turbofan,turbojets, turboshafts, and three-spool (plus fan) turbofans wherein anintermediate spool includes an intermediate pressure compressor (“IPC”)between a Low Pressure Compressor (“LPC”) and a High Pressure Compressor(“HPC”), and an intermediate pressure turbine (“IPT”) between the highpressure turbine (“HPT”) and the Low pressure Turbine (“LPT”).

The engine 20 generally includes a low spool 30 and a high spool 32mounted for rotation about an engine central longitudinal axis Arelative to an engine static structure 36 via several bearingcompartments 38. The low spool 30 generally includes an inner shaft 40that interconnects a fan 42, a low pressure compressor 44 (“LPC”) and alow pressure turbine 46 (“LPT”). The inner shaft 40 drives the fan 42directly or thru a geared architecture 48 to drive the fan 42 at a lowerspeed than the low spool 30. An exemplary reduction transmission is anepicyclic transmission, namely a planetary or star gear system.

The high spool 32 includes an outer shaft 50 that interconnects a highpressure compressor 52 (“HPC”) and high pressure turbine 54 (“HPT”). Acombustor 56 is arranged between the HPC 52 and the HPT 54. The innershaft 40 and the outer shaft 50 are concentric and rotate about theengine central longitudinal axis A which is collinear with theirlongitudinal axes.

Core airflow is compressed by the LPC 44 then the HPC 52, mixed withfuel and burned in the combustor 56, then expanded over the HPT 54 andthe LPT 46. The turbines 54, 46 rotationally drive the respective lowspool 30 and high spool 32 in response to the expansion. The main engineshafts 40, 50 are supported at a plurality of points by the bearingcompartments 38. It should be understood that various bearingcompartments 38 at various locations may alternatively or additionallybe provided.

In one example, the gas turbine engine 20 is a high-bypass gearedaircraft engine with a bypass ratio greater than about six (6:1). Thegeared architecture 48 can include an epicyclic gear train, such as aplanetary gear system or other gear system. The example epicyclic geartrain has a gear reduction ratio of greater than about 2.3:1, and inanother example is greater than about 2.5:1. The geared turbofan enablesoperation of the low spool 30 at higher speeds which can increase theoperational efficiency of the LPC 44 and LPT 46 to render increasedpressure in a relatively few number of stages.

A pressure ratio associated with the LPT 46 is pressure measured priorto the inlet of the LPT 46 as related to the pressure at the outlet ofthe LPT 46 prior to an exhaust nozzle of the gas turbine engine 20. Inone non-limiting embodiment, the bypass ratio of the gas turbine engine20 is greater than about ten (10:1), the fan diameter is significantlylarger than that of the LPC 44, and the LPT 46 has a pressure ratio thatis greater than about five (5:1). It should be understood, however, thatthe above parameters are only exemplary of one embodiment of a gearedarchitecture engine and that the present disclosure is applicable toother gas turbine engines including direct drive turbofans, where therotational speed of the fan 42 is the same (1:1) of the LPC 44.

In one example, a significant amount of thrust is provided by the bypassflow path due to the high bypass ratio. The fan section 22 of the gasturbine engine 20 is designed for a particular flightcondition—typically cruise at about 0.8 Mach and about 35,000 feet. Thisflight condition, with the gas turbine engine 20 at its best fuelconsumption, is also known as bucket cruise Thrust Specific FuelConsumption (TSFC). TSFC is an industry standard parameter of fuelconsumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fansection 22 without the use of a Fan Exit Guide Vane system. Therelatively low Fan Pressure Ratio according to one example gas turbineengine 20 is less than 1.45. Low Corrected Fan Tip Speed is the actualfan tip speed divided by an industry standard temperature correction of(“T”/518.7)^(0.5) in which “T” represents the ambient temperature indegrees Rankine. The Low Corrected Fan Tip Speed according to oneexample gas turbine engine 20 is less than about 1150 fps (351 m/s).

With reference to FIG. 2, the combustor 56 generally includes an outercombustor liner assembly 60, an inner combustor liner assembly 62 and adiffuser case module 64. The outer combustor liner assembly 60 and theinner combustor liner assembly 62 are spaced apart such that acombustion chamber 66 is defined there between. The combustion chamber66 may be generally annular in shape.

The outer combustor liner assembly 60 is spaced radially inward from anouter diffuser case 64A of the diffuser case module 64 to define anouter annular plenum 76. The inner combustor liner assembly 62 is spacedradially outward from an inner diffuser case 64B of the diffuser casemodule 64 to define an inner annular plenum 78. It should be understoodthat although a particular combustor is illustrated, other combustortypes with various combustor liner arrangements will also benefitherefrom. It should be further understood that the disclosed coolingflow paths are but an illustrated embodiment and should not be limitedonly thereto.

The combustor liner assemblies 60, 62 contain the combustion productsfor direction toward the turbine section 28. Each combustor linerassembly 60, 62 generally includes a respective support shell 68, 70which supports one or more heat shields 72, 74 mounted to a hot side ofthe respective support shell 68, 70. Each of the heat shields 72, 74 maybe generally rectilinear and manufactured of, for example, a nickelbased super alloy, ceramic or other temperature resistant material andare arranged to form a liner array. In one disclosed non-limitingembodiment, the liner array includes a multiple of forward heat shields72A and a multiple of aft heat shields 72B that are circumferentiallystaggered to line the hot side of the outer shell 68. A multiple offorward heat shields 74A and a multiple of aft heat shields 74B arecircumferentially staggered to line the hot side of the inner shell 70.

The combustor 56 further includes a forward assembly 80 immediatelydownstream of the compressor section 24 to receive compressed airflowtherefrom. The forward assembly 80 generally includes an annular hood82, a bulkhead assembly 84, a multiple of fuel nozzles 86 (one shown)and a multiple of pre-swirlers 90 (one shown). Each of the pre-swirlers90 is circumferentially aligned with one of a respective annular hoodport 94 and projects thru the bulkhead assembly 84. The bulkheadassembly 84 generally includes a bulkhead support shell 96 secured tothe combustor liner assembly 60, 62, and a multiple of circumferentiallydistributed bulkhead heat shields 98 secured to the bulkhead supportshell 96 to define an opening 92 for each pre-swirler 90.

The annular hood 82 extends radially between, and is secured to, theforwardmost ends of the combustor liner assemblies 60, 62. Each fuelnozzle 86 may be secured to the diffuser case module 64 and project thruone of the hood ports 94 and the respective pre-swirler 90. Each of themultiple of circumferentially distributed hood ports 94 accommodates therespective fuel nozzle 86 to introduce air into the forward end of thecombustion chamber 66 thru the pre-swirler 90.

The forward assembly 80 introduces core combustion air into the forwardsection of the combustion chamber 66 while the remainder enters theouter annular plenum 76 and the inner annular plenum 78. The multiple offuel nozzles 86 and adjacent structure generate a blended fuel-airmixture that supports stable combustion in the combustion chamber 66.

Opposite the forward assembly 80, the outer and inner support shells 68,70 are mounted to a first row of Nozzle Guide Vanes (NGVs) 54A in theHPT 54. The NGVs 54A are static components that direct the combustiongases onto the turbine blades of the first turbine rotor in the turbinesection 28 to facilitate the conversion of pressure energy into kineticenergy. The combustion gases are also accelerated by the NGVs 54Abecause of their convergent shape and are typically given a “spin” or a“swirl” about axis A in the direction of turbine sections' rotation.

The inner diffuser case 64B defines an annular flow path 100 forcompressed airflow C from the upstream HPC 52. The annular flow path 100includes a multiple of struts 102 which extend in a radial directionbetween an outer shroud 104 and an inner shroud 106 (also shown in FIG.3). The annular flow path 100 defines a flowpath temperature profile(illustrated schematically at 110; FIG. 4) at the exit of the HPC 52that is non-uniform, with a relatively cooler mid-span pre-diffuserairflow with respect to a relatively hotter outer diameter airflowadjacent to the outer shroud 104 and a relatively hotter inner diameterairflow adjacent to the inner shroud 106. It should be understood thatthe relatively longer arrows of the flow path temperature profile 110correspond to relatively higher temperatures.

With reference to FIG. 4, the radially non-uniform airflow temperatureprofile 110 typically communicates with the combustor 56, however,increased turbine section 28 durability and/or the ability to withstandhotter turbine section 28 flowpath temperatures are readily achievedwhen the relatively cooler pre-diffuser mid-span airflow is directlytapped for use in other engine sections.

To tap the relatively cooler pre-diffuser airflow, a mid-spanpre-diffuser inlet 112 is located between the shrouds 104, 106 generallyparallel to the core airflow from the HPC 52 to collect and duct therelatively cooler mid-span airflow to desired regions Z within theengine 20 thru a manifold 114 (illustrated schematically; FIG. 5). Thepre-diffusor inlet 112 may be parallel to an airflow from the HPC 52 andnot necessarily parallel to the shrouds 104, 106. That is, thepre-diffuser inlet 112 may be oriented with respect to the airflow andnot the associated hardware. As defined herein, mid-span is any radiallocation between the shrouds 104, 106 and is not to be limited to onlythe exact middle of the struts 102. The pre-diffuser inlet 112 may beselectively located radially along the strut 102 to essentially selectfrom the airflow temperature profile 110. That is, the radial positionmay be predefined to tap a temperature tailored airflow. Oftentimes,even a relatively small temperature differential provides advantageoususage in other regions Z.

The manifold 114 as well as those that are hereafter described may be ofvarious constructions and geometries to include but not limited toconduits as well as integral passageways within engine static structuresuch as the diffuser case 64. Furthermore, directional structures suchas turning vanes and other guides may also be incorporated in themanifold 114 to minimize flow loss.

With reference to FIG. 5, the temperature tailored airflow tapped fromthe mid-span pre-diffuser inlet 112 may be communicated to the variousregions Z within the engine 20 such as, for example only, the HPC 52,the combustor 56, the HPT 54 bearing compartments 38, or other enginearchitecture sections such as an exhaust duct section, an augmentersection, roll posts or other regions. That is, the non-uniform airflowtemperature 110 is selectively tapped by radial location of the mid-spanpre-diffuser inlet 112 to provide a temperature tailored airflow 111(illustrated schematically) for communication to the desired region Z.The temperature tailored airflow 111 is thereby tailored in response tothe radial tap position of the mid-span pre-diffuser inlet 112 along thespan of the struts 102. Furthermore, the temperature tailored airflow111 may be selectively communicated thru a heat exchanger to stillfurther modify the temperature tailored airflow 111. It should beappreciated that all or only a portion of the temperature tailoredairflow tapped from the mid-span pre-diffuser inlet 112 may becommunicated thru the heat exchanger.

With reference to FIG. 6, the struts 102 are defined by an outer airfoilwall surface 116 between a leading edge 118 and a trailing edge 120. Theouter airfoil wall surface 116 may define a generally concave shapedportion to form a pressure side 122 and a generally convex shapedportion forming a suction side 124 (FIG. 7). It should be appreciatedthat various airfoil and non-airfoil shapes may alternatively beprovided.

The mid-span pre-diffuser inlet 112 according to one disclosednon-limiting embodiment may include a flush wall NACA inlet 126 oneither or both sides 122, 124 of one or more of the struts 102 (FIG. 7).That is, the flush wall NACA inlet 126 is located at least partiallywithin the outer airfoil wall surface 116. A capacity of approximately0%-2.5% of the airflow from the HPC 52 may be typically provided by eachNACA inlet 126.

With reference to FIG. 8, the mid-span pre-diffuser inlet 112 accordingto another disclosed non-limiting embodiment may include an enhancedcapacity side-winged mid-span pre-diffuser RAM inlet 128 that extendsfrom one or both sides of one or more struts 102′ (FIG. 9). That is, theside-winged mid-span pre-diffuser RAM inlet 128 extends outward from theouter airfoil wall surface 116 to receive RAM airflow. A capacity ofapproximately 2.5%-5% of the airflow from the HPC 52 may be typicallyprovided by each RAM inlet 128.

With reference to FIG. 10, the mid-span pre-diffuser inlet 112 accordingto another disclosed non-limiting embodiment may include an annularinlet 130 (also shown in FIG. 11). The annular inlet 130 is locatedcircumferentially between the inlet struts 102 and radially between theouter and inner shrouds 104, 106. That is, the annular inlet 130 extendsbetween the outer airfoil wall surface 116 of adjacent inlet struts 102in a circumferentially segmented arrangement. Alternatively, the annularinlet 130 according to another disclosed non-limiting embodiment mayextend between and at least partially thru the outer airfoil wallsurface 116 in a substantially circumferentially continuous arrangement.A capacity of approximately 10%-20% of the airflow from the HPC 52 maybe typically provided by annular inlet 130.

With reference to FIG. 12, the relatively hotter airflow from the outerdiameter zone adjacent to the outer shroud 104 and the relatively hotterairflow from the inner diameter zone adjacent to the inner shroud 106 isdirected by a pre-swirler manifold 132 (illustrated schematically byarrows) to the combustor fuel-nozzle pre-swirlers 90. A capacity ofapproximately 80% of the airflow from the HPC 52 may be provided hereby.Furthermore, by tapping the mid-span pre-diffuser airflow with themid-span pre-diffuser inlets 112, the average temperature of the airflowprovided to the fuel-nozzle pre-swirlers 90 even without the manifold132 is relatively higher.

Provision of relatively hotter endwall air to the combustor fuel-nozzlepre-swirlers 90 facilitates a performance benefit as less fuel isrequired to heat the core airflow to a target level within thecombustion chamber 66 referred to herein as T4. As further perspective,T1 is a temperature in front of the fan section 22; T2 is a temperatureat the leading edge of the fan 42; T2.5 is the temperature between theLPC 44 and the HPC 52; T3 is the temperature aft of the LPC 44; T4 isthe temperature in the combustion chamber 66; T4.5 is the temperaturebetween the HPT 54 and the LPT 46; and T5 is the temperature aft of theLPT 46 (FIG. 1). The relatively hotter endwall air provided to thecombustor fuel-nozzle pre-swirlers 90 may for example, provide anapproximately 50° F. (10° C.) performance benefit.

With reference to FIG. 13, according to another disclosed non-limitingembodiment, the relatively cooler mid-span airflow is communicated via amanifold 134 (illustrated schematically by an arrow) to HPC aft rotorblade attachments 136. The rear hub 138 may, in part, define themanifold 134. It should be appreciated, however, that various structuresand airflow paths may alternatively or additionally be provided.Approximately 1%-1.5% of the airflow from the HPC 52 may be typicallyprovided thru the manifold 134 to cool and purge the HPC aft rotor bladeattachments 136.

With reference to FIG. 14, according to another disclosed non-limitingembodiment, the relatively cooler mid-span airflow is first communicatedthru a heat exchanger 140 prior to communication thru the manifold 134.The heat exchanger 140 further lowers the air temperature of the airflowfrom the HPC 52 which facilitates additional increases in turbinedurability and/or gaspath temperature capability. It should beappreciated that the heat exchanger 140, and those that follow, may beselectively operable and located in various sections of the engine 20.

With reference to FIG. 15, according to another disclosed non-limitingembodiment, the airflow from the heat exchanger 140 may also becommunicated as buffer air thru a buffer passage 142 (illustratedschematically by an arrow) to one or more bearing compartments 38. Thebuffer air maintains a positive differential pressure across seals inthe bearing compartment 38 to facilitate a pneumatic pressure barrier toprevent undesired oil leakage therefrom. Approximately 0.25%-0.5% of theairflow from the HPC 52 may be provided as buffer air that is typicallyat temperatures of approximately 450° F. (232° C.). In one disclosed,non-limiting embodiment, the buffer passage 142 may pass spanwise thruone or more struts 102B (FIG. 16). It should be appreciated that anynumber of struts 102B will benefit herefrom. The struts 102B with thespanwise passage may or may not include mid-span pre-diffuser inlets112.

With reference to FIG. 17, according to another disclosed non-limitingembodiment, the relatively cooler mid-span airflow is communicated via amanifold 144 (illustrated schematically by an arrow) to an HPC aft rotorhub 146. The HPC aft rotor hub 146 may, in part, define the manifold144. It should be appreciated, however, that various structures andairflow paths may alternatively or additionally be provided.

In another disclosed non-limiting embodiment, the relatively coolermid-span airflow is first communicated thru a heat exchanger 148 priorto communication thru the manifold 144 to further lower the airtemperature of the airflow from the HPC 52 which facilitates additionalincreases in turbine durability and/or gaspath temperature capability.

In another disclosed non-limiting embodiment, the airflow from the heatexchanger 148 may also be communicated as buffer air thru a bufferpassage 150 (illustrated schematically by an arrow) to one or morebearing compartments 38. Although the disclosed illustrated embodimentsare directed to the final stages of the HPC 52, it should be appreciatedthat any number of stages will benefit herefrom as well as other enginesections such as the LPC 44 and other bearing compartments.

With reference to FIG. 18, according to another disclosed non-limitingembodiment, the relatively cooler mid-span airflow is communicated via amanifold 152 (illustrated schematically by arrows) to the combustor 56to tailor the burner exit temperature profile 154 (illustratedschematically). That is, by selective direction and communication of therelatively cooler mid-span airflow to the combustor 56. For example, thequench flow typical of a Rich-Quench-Lean type combustor may be tailoredsuch that the burner exit temperature profile 154 is the inverse of theairflow from the HPC 52. That is, the burner exit temperature profile154 is non-uniform, with a relatively hotter annular mid-span zone withrespect to a relatively cooler outer diameter zone adjacent to an outershroud 156 of the NGVs 54A and a relatively hotter inner diameter zoneadjacent to an inner shroud 158 of the NGVs 54A. It should beappreciated that various combustor types and exit temperature profileswill also benefit herefrom.

In another disclosed non-limiting embodiment, the relatively coolermid-span airflow is first communicated thru a heat exchanger 160 priorto communication thru the manifold 152 to further lower the airtemperature of the airflow from the HPC 52 which facilitates furthercontrol of the burner exit temperature profile 154.

In another disclosed non-limiting embodiment, the airflow from the heatexchanger 160 may also be communicated as buffer air thru a bufferpassage 162 (illustrated schematically by an arrow) to one or morebearing compartments 38.

With reference to FIG. 19, according to another disclosed non-limitingembodiment, the relatively cooler mid-span airflow is communicated via amanifold 164 (illustrated schematically by arrows) to the NGVs 54A whichare also referred to as the 1st stage vanes of the HPT 54. Therelatively cooler mid-span airflow may supply or otherwise supplement asecondary airflow to the NGVs 54A. That is, a mixed flow may beprovided.

In another disclosed non-limiting embodiment, the relatively coolermid-span airflow is first communicated thru a heat exchanger 166 priorto communication thru the manifold 164 to further lower the airtemperature of the airflow from the HPC 52 which facilitates additionalincreases in turbine durability and/or gaspath temperature capability.

In another disclosed non-limiting embodiment, the airflow from the heatexchanger 166 may also be communicated as buffer air thru a bufferpassage 168 (illustrated schematically by an arrow) to one or morebearing compartments 38.

With reference to FIG. 20, according to another disclosed non-limitingembodiment, the relatively cooler mid-span airflow is communicated via amanifold 170 (illustrated schematically by an arrow) to a 1st stageblade outer air seal (BOAS) 172 of the HPT 54. The BOAS may beinternally cooled with cooling air communicated into an outboard plenumof the BOAS 172 then pass thru passageways in the seal body and exitoutlet ports thru an inboard side to provide film cooling. Therelatively cooler mid-span airflow may also exit along thecircumferential matefaces of the BOAS 172 so as to be vented into anadjacent inter-segment region to, for example, cool feather seals theadjacent BOAS segments.

In another disclosed non-limiting embodiment, the relatively coolermid-span airflow is first communicated thru a heat exchanger 174 priorto communication thru the manifold 170 to further lower the airtemperature of the airflow from the HPC 52 which facilitates additionalincreases in turbine durability and/or gaspath temperature capability.

In another disclosed non-limiting embodiment, the airflow from the heatexchanger 174 may also be communicated as buffer air thru a bufferpassage 176 (illustrated schematically by an arrow) to one or morebearing compartments 38.

With reference to FIG. 21, according to another disclosed non-limitingembodiment, the relatively cooler mid-span airflow is communicated via amanifold 178 (illustrated schematically by arrows) to 1st stage blades180 of the HPT 54. The manifold 178 may communicate with a bladeattachment region 182 of a rotor disk 184 that supports the blades 180.That is, the manifold 178 may communicate thru a rotor disk cover plate186 to direct the relatively cooler mid-span airflow into the bladeattachment region 182 and thence into the blades 180 thru respectiveroot section thereof.

Furthermore, the manifold 178 may communicate with the rotor disk coverplate 186 thru a tangential on board injector (TOBI) 188. The TOBI 188is often known by other names but generally includes annular spacednozzles that impart a swirling moment to direct the airflowtangentially.

In another disclosed non-limiting embodiment, the relatively coolermid-span airflow is first communicated thru a heat exchanger 190 priorto communication thru the manifold 178 to further lower the airtemperature of the airflow from the HPC 52 which facilitates additionalincreases in turbine durability and/or gaspath temperature capability.

In another disclosed non-limiting embodiment, the airflow from the heatexchanger 190 may also be communicated as buffer air thru a bufferpassage 192 (illustrated schematically by an arrow) to one or morebearing compartments 38.

With reference to FIG. 22, according to another disclosed non-limitingembodiment, the relatively cooler mid-span airflow is communicated via amanifold 194 (illustrated schematically by arrows) to 2nd stage vanes196 of the HPT 54. The relatively cooler mid-span airflow may supply orotherwise supplement a secondary airflow to the 2nd stage vanes 196.Although the disclosed illustrated embodiments are directed to atwo-stage HPT 54, it should be appreciated that any number of stageswill benefit herefrom as well as other engine sections such as the LPT46.

In another disclosed non-limiting embodiment, the relatively coolermid-span airflow is first communicated thru a heat exchanger 198 priorto communication thru the manifold 194 to further lower the airtemperature of the airflow from the HPC 52 which facilitates additionalincreases in turbine durability and/or gaspath temperature capability.

In another disclosed non-limiting embodiment, the airflow from the heatexchanger 198 may also be communicated as buffer air thru a bufferpassage 200 (illustrated schematically by an arrow) to one or morebearing compartments 38.

With reference to FIG. 23, according to another disclosed non-limitingembodiment, the relatively cooler mid-span airflow is communicated via amanifold 202 (illustrated schematically by arrows) to the 1-2 seals 204of the HPT 54. The relatively cooler mid-span airflow may supply orotherwise supplement a secondary airflow.

In another disclosed non-limiting embodiment, the relatively coolermid-span airflow is first communicated thru a heat exchanger 206 priorto communication thru the manifold 202 to further lower the airtemperature of the airflow from the HPC 52 which facilitates additionalincreases in turbine durability and/or gaspath temperature capability.

In another disclosed non-limiting embodiment, the airflow from the heatexchanger 206 may also be communicated as buffer air thru a bufferpassage 208 (illustrated schematically by an arrow) to one or morebearing compartments 38.

Although the different non-limiting embodiments have specificillustrated components, the embodiments of this invention are notlimited to those particular combinations. It is possible to use some ofthe components or features from any of the non-limiting embodiments incombination with features or components from any of the othernon-limiting embodiments.

It should be understood that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like arewith reference to the normal operational attitude of the vehicle andshould not be considered otherwise limiting.

It should be understood that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be understood that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom.

Although particular step sequences are shown, described, and claimed, itshould be understood that steps may be performed in any order, separatedor combined unless otherwise indicated and will still benefit from thepresent disclosure.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beunderstood that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

What is claimed is:
 1. A gas turbine engine comprising: a compressor section; a combustor section; a diffuser case module comprising an inner shroud, an outer shroud and a multiple of struts within an annular flow path from said compressor section to said combustor section, said multiple of struts comprising a first strut that extends radially between said inner shroud and said outer shroud, wherein said first strut includes a leading edge, a trailing edge and an outer wall surface extending between the leading edge and the trailing edge, wherein said first strut defines a mid-span pre-diffuser inlet in communication with said annular flow path, and wherein said mid-span pre-diffuser inlet is on a side of the first strut and in the outer wall surface; and a manifold in communication with said mid-span pre-diffuser inlet and said compressor section; wherein said first strut has a circumferential width that increases as said first strut extends from said leading edge to said trailing edge such that said circumferential width has a first value at said leading edge that is less than a second value of said circumferential width at said trailing edge.
 2. The gas turbine engine of claim 1, wherein said mid-span pre-diffuser inlet comprises a NACA inlet.
 3. A gas turbine engine comprising: a compressor section; a combustor section; a diffuser case module comprising an inner shroud, an outer shroud and a multiple of struts within an annular flow path from said compressor section to said combustor section, said multiple of struts comprising a first strut that extends radially between said inner shroud and said outer shroud, wherein said first strut includes a leading edge, a trailing edge and an outer wall surface extending between the leading edge and the trailing edge, wherein said first strut defines a mid-span pre-diffuser inlet in communication with said annular flow path, and wherein said mid-span pre-diffuser inlet is on a side of the first strut and in the outer wall surface; and a manifold in communication with said mid-span pre-diffuser inlet and said compressor section; wherein said first strut has a maximum circumferential width at a first location along a longitudinal length of said first strut, and said first location is at said trailing edge; and wherein said mid-span pre-diffuser inlet is located at a second location along said longitudinal length of said first strut where said first strut has a second circumferential width that is less than said maximum circumferential width.
 4. A gas turbine engine comprising: a compressor section; a combustor section; a diffuser case module comprising an inner shroud, an outer shroud and a multiple of struts within an annular flow path from said compressor section to said combustor section, said multiple of struts comprising a first strut that extends radially between said inner shroud and said outer shroud, wherein said first strut includes a leading edge, a trailing edge and an outer wall surface extending between the leading edge and the trailing edge, wherein said first strut defines a mid-span pre-diffuser inlet in communication with said annular flow path, and wherein said mid-span pre-diffuser inlet is on a side of the first strut and in the outer wall surface; and a manifold in communication with said mid-span pre-diffuser inlet and said compressor section such that said compressor section is fluidly coupled with said mid-span pre-diffuser inlet through said manifold; wherein said mid-span pre-diffuser inlet comprises a scoop that projects into said annular flow path from said side of said first strut; wherein said first strut has a maximum circumferential width at a first location along a longitudinal length of said first strut, and said first location is at said trailing edge; and wherein said mid-span pre-diffuser inlet is located at a second location along said longitudinal length of said first strut where said first strut has a second circumferential width that is less than said maximum circumferential width.
 5. The gas turbine engine as recited in claim 4, wherein said manifold communicates a temperature tailored airflow.
 6. The gas turbine engine as recited in claim 4, wherein said manifold communicates a temperature tailored airflow through a heat exchanger.
 7. The gas turbine engine as recited in claim 6, wherein said manifold communicates said temperature tailored airflow from said heat exchanger as buffer air.
 8. The gas turbine engine as recited in claim 7, wherein said buffer air is communicated through a buffer passage to one or more bearing compartments.
 9. The gas turbine engine as recited in claim 4, wherein said mid-span pre-diffuser inlet supplies a temperature tailored airflow into said manifold.
 10. The gas turbine engine as recited in claim 4, wherein said manifold communicates with a high pressure compressor of said compressor section.
 11. The gas turbine engine as recited in claim 4, wherein said manifold is annular.
 12. The gas turbine engine as recited in claim 4, wherein said manifold communicates with a row of rotor blade attachments in said compressor section.
 13. The gas turbine engine as recited in claim 12, wherein said manifold is at least partially defined by a rotor hub.
 14. The gas turbine engine of claim 4, wherein said scoop is a side-winged RAM inlet. 